1. Field of the Invention
The present invention relates to an apparatus and method for enhancing flow mixing in the boundary layer in the inlet of a turbine engine for aircraft operating at supersonic velocities, and more particularly, to apparatus and methods for inhibiting flow separation from the wall of a supersonic engine inlet without undue penalties in engine performance.
2. Description of Related Art
Supersonic aircraft present numerous engineering challenges due to the shock waves they generate when operating at Mach numbers greater than one (M>1). A particularly difficult area to properly engineer is the aircraft turbine engine inlet diffuser. A large number of these challenges result from the necessity of slowing the supersonic flow entering the engine diffuser to a subsonic velocity (M<1) before it enters the aircraft engine compressor. It is axiomatic that slowing from supersonic to subsonic flow requires as a practical matter that the flow pass through one or more shock waves. This introduces irreversible total pressure losses, which can be mitigated to some extent by optimizing the shock structure in a supersonic inlet diffuser.
The large majority of supersonic engine inlets assume one of a number of basic configurations that have been settled for some time, and these configurations generate fairly well understood shock structures, as discussed in more detail further below. A standard inlet configuration seeks to minimize the inevitable pressure losses encountered in decelerating from supersonic to subsonic Mach numbers. The exact inlet dimensions and configuration must be determined for each new aircraft, and depend on numerous factors, such as the aircraft's overall mission profile, the airflow requirements of the engine at different points in the mission profile, and many others. This can present a complex design problem and sometimes requires that provision be made for altering the inlet configuration during aircraft operation. Nevertheless, the basic principles are well known and sophisticated design tools such as computational fluid dynamics are available to the engine and aircraft design engineer for optimizing inlet configurations for particular aircraft.
However, the presence of shock waves in the flow also causes other less well understood phenomena. One of these is the shock wave/boundary layer interaction (SBLI) within the diffuser's inlet duct. A boundary layer forms in a fluid moving relative to a wall due to viscous properties of the fluid, because the flow velocity is zero at the wall surface but accelerates to the free stream flow velocity at some point spaced from the wall. The distance from the wall to where the flow attains the free stream velocity is the boundary layer thickness, which generally tends to increase from one location to another traveling downstream in the fluid flow adjacent to a wall. The interaction of the boundary layer with the shock in a supersonic inlet diffuser has been the subject of many technical articles and patents. One such article is Herges, T., et al., “Micro-Ramp Flow Control on Normal Shock/Boundary Later Interactions,” Amer. Inst. of Aeronautics and Astronautics Rept. No. AIAA 2009-920 (January 2009) (“Herges”). As Herges points out, in a typical supersonic inlet configuration the air flow decelerates through a series of shock structures usually terminating in a normal shock within an inlet duct. As a result, the growing boundary layer along the inlet duct wall experiences adverse pressure gradients, possibly resulting in boundary layer separation and unsteady pressure fluctuations in the pressure field downstream of the normal shock that adversely affect engine compressor performance and can even cause engine un-start. Herges notes that bleed systems, which extract low-momentum fluid from the boundary layer, have been demonstrated to reduce both boundary layer thickness and the severity of separation. However, bleeding fluid from the inlet flow results in a decrease of mass flow of air to the engine. Since a particular engine design must satisfy the thrust requirements necessary for the aircraft to operate throughout its mission profile, and engine thrust depends in significant part on the mass flow of air through the engine, a reduction in mass flow from bleeding the boundary layer usually requires a larger inlet to compensate for the reduction. This in turn causes more drag, adds weight (thus reducing aircraft payload), and lowers engine efficiency.
Herges discusses several flow-control methods that have been proposed in addition to boundary layer bleed systems to control SBLI. Early investigation into SBLI control focused on two-dimensional surface features that aimed to spread out the bifurcated lambda foot of the normal shock. A lambda shock decelerates the flow through a pair of oblique shocks, thus having a smaller total pressure loss than a single normal shock with the same static pressure rise. In addition, by spreading the lambda shock, the boundary layer will experience a smaller adverse pressure gradient and will be less likely to separate. This approach to SBLI control has been investigated by Herges and others using a variety of methodologies including porous surfaces and two-dimensional bumps. While two-dimensional surface features like bumps proved effectual at spreading the lambda foot of the normal shock, such large surface features increase drag.
Herges indicates that investigations into SBLI control using arrays of three-dimensional geometries for SBLI control suggested using sub-boundary layer vortex generators (SBVGs) for reducing boundary layer growth and eliminating or reducing separation when placed upstream of the shock wave/boundary layer interaction. In contrast with other approaches to SBLI control, SBVGs aim to control SBLI by energizing the boundary layer and making it less susceptible to shock-induced separation. SBVGs entrain higher momentum fluid to energize the low momentum fluid near the wall and improve boundary layer health and suppress or delay separation. Due to their smaller size, SBVGs have been shown to have significantly reduced device drag compared to conventional vortex generators.
One SBVG geometry that has been of particular interest is the micro-ramp. Herges states that micro-ramps can reduce separation region length and boundary layer thickness while being physically robust and easy to machine. While micro-ramps did not completely eliminate boundary layer separation in all experiments, they have been shown to be effective in breaking up separation regions in the vicinity directly behind the micro-ramp. Therefore, micro-ramps optimally would be provided in an array in the engine inlet. Herges presents the results of experiments exploring possible geometries of micro-ramps and how to arrange them in arrays for optimal results.
There are still other technical articles reporting the results of investigations into the use of so-called micro-ramps or other configurations of SBVGs. One of these papers is Anderson, B., et al., “Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation,” Amer. Inst. of Aeronautics and Astronautics Rept. No. AIAA 2006-3197 (June 2006) (“Anderson”). One of the objectives of Anderson's reported experiments was to establish the ability of micro-array flow control to manage the shock wave/turbulent boundary layer interactions in supersonic inlets and to evaluate the effectiveness of micro-array flow control relative to conventional inlet boundary layer bleed. Anderson notes that the purpose of the micro-arrays was to alter the properties of the supersonic boundary layer by introducing a cascade of counter-rotating micro-vortices in the near wall region surrounding the inlet surfaces. Anderson concludes that “the impact of the SWBL [shock wave boundary layer] interaction on the main flow field was minimized without boundary bleed.”
Thus, the art has long recognized the efficacy of small, ramp-like, vortex-generating structures in minimizing the impact on supersonic engine inlet flow of shock wave boundary layer interaction, thereby eliminating the need for boundary layer bleed and its concomitant disadvantages. In addition to the Herges and Anderson articles, the patent literature includes numerous examples of different micro-ramp geometries and array configurations that act as vortex generators in a variety of environments. Some examples are shown in U.S. Pat. Nos. 3,578,264, 4,455,045, 4,175,640, 4,487,017, 4,696,442, 5,058,837, 5,598,990 and 8,087,250, and Pub. No. US 2010/0288379.
However, the fixed-geometry, permanently deployed vortex generators in the known prior art must inherently be designed for maximum relief of shock wave boundary layer interactions at a single point in the engine's operating envelope. Whatever design point is chosen (say, level-cruise at altitude), the presence of the vortex generators can actually degrade engine performance at other important engine operating points (such as climb under power). Any performance penalty is highly undesirable, and it is expected that adverse effects on engine efficiency will be particularly closely scrutinized in possible future commercial supersonic aircraft, since commercial aircraft operators generally consider efficiency penalties even in fractions of one percent to be unacceptable. In addition, a plurality of the prior art micro-ramps arrayed along a substantial portion of a wall of a supersonic engine inlet cause increased drag and concomitant pressure losses. These must be made up in some fashion if the engine is still to meet the performance requirements dictated by the aircraft mission. Typically, this involves increasing the inlet size to increase the mass flow of air through the engine to make up for the lost performance. In this respect, micro-ramps exhibit to an extent the undesirable side effects of using boundary layer bleed.
Further, it was noted above that the interaction of the boundary layer and the normal shock in the inlet also causes unsteady flow phenomena, and the dynamic behavior of the shock structures in the engine inlet can also have a significant adverse impact on the stability of the flow through the engine compressor. Flow instabilities due to SBLI can cause engine performance problems such as compressor stall and surge and a compressor blade aeroelastic effect known as “compressor buzz.” While there has been extensive research aimed at further understanding, and overcoming, many adverse effects of non-steady-state flow in supersonic inlets, such research is ongoing and can be expected to continue to suggest new ways of mitigating such effects. A fixed configuration array of micro-ramps or other types of fixed-geometry vortex generators will be inherently limited in their ability to adapt to new inlet configurations that might be adopted for a given engine or aircraft after its initial design parameters are set. Fixed configuration, fixed geometry vortex generators are also unable to adapt to changes in the mission profile of a particular aircraft or to changes in engine design after the inlet has been fabricated with a fixed vortex generator configuration and geometry.
It would be advantageous if known micro-ramps or other vortex generators had a variable configuration, for example, by being retractable out of the air flow in the inlet during flight regimes when they are not needed and deployable when desired. In addition, not all flight regimes require the same amount of vorticity to be introduced into the flow in order to mitigate the effects of shock wave/boundary layer interaction. Accordingly, it would be likewise advantageous to be able to control the vortex generator geometry to account for different conditions associated with different flight regimes. Most of the prior art vortex generators are solid ramps, such as those used in the Herges and Anderson experiments discussed above, or other configurations that would be equally difficult to construct for deployment into and retraction from the flow, or for changing their geometry to control the characteristics of the generated vortices. And even if the advantages of providing deployable and retractable vortex generators had been recognized, mechanical linkages and complex mounting configurations might have been envisioned for prior art micro-ramps or other vortex generator structures. But it is not known that anyone before now even contemplated making prior art micro-ramps or other vortex generating structures for supersonic engine inlets retractable or capable of multiple positions to vary the properties of the vortices they generate.
There is at least one deployable vortex generator structure known in the prior art, as disclosed in U.S. Pat. No. 7,798,448, assigned to the assignee of the present invention. This patent discloses a flow-driven oscillating flap that attenuates noise generated by flow over a cavity. A deploying mechanism uses a shape-memory alloy (SMA) to provide a motive force to deploy the flap into the flow and retract it out of the flow. However, nothing in this patent suggests deployable/retractable vortex generators for supersonic engine inlets or providing for deployment through a range of positions to control the properties of the generated vortices. In fact, the assignee of the present invention is a pioneer in the use of shape-memory alloys to actuate flow control surfaces, as evidenced by others of its patents, such as U.S. Pat. Nos. 5,752,672, 6,220,550, and 6,497,385, but even so has not before now turned its attention to using shape memory alloy actuation to provide deployable, retractable, and controllable vortex generators for a supersonic aircraft engine inlet diffuser.